Present embodiments relate generally to apparatus, methods, and/or systems for improving durability of a nozzle segment for a turbine engine. More specifically, not by way of limitation, present embodiments relate to a more durable vane for a nozzle of a gas turbine engine having improved cooling capacity and allowing for higher temperature operation without negatively impacting aero-performance.
A gas turbine engine generally includes a compressor, a combustor and a turbine. The compressor and turbine generally include rows of airfoils that are stacked axially in stages. Each stage includes a row of circumferentially spaced stator vanes and a row of rotor blades which rotate about a center shaft or axis of the turbine engine.
In the turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages. These turbine stages extract energy from the combustion gases. A high pressure turbine includes a first stage nozzle and a rotor assembly including a disk and a plurality of turbine blades. The high pressure turbine first receives the hot combustion gases from the combustor and includes a first stage stator nozzle that directs the combustion gases exiting from the combustor downstream through a row of high pressure turbine rotor blades extending radially outwardly from a first rotor disk. For embodiments including two or more turbine stages, each stage comprises a stator nozzle with a set of rotor blades. For example, not meant to be limiting, of a two stage turbine, a second stage stator nozzle is positioned downstream of the first stage blades followed in turn by a row of second stage turbine blades extending radially outwardly from a second rotor disk. The stator nozzles direct and aim the hot combustion gas in a manner to maximize extraction at the adjacent downstream turbine blades.
The first and second rotor disks are joined to the compressor by a corresponding rotor shaft for powering the compressor during operation. The turbine engine may include a number of stages of static airfoils, commonly referred to as vanes, interspaced in the engine axial direction between rotating airfoils commonly referred to as blades. A multi-stage low pressure turbine follows the two stage high pressure turbine and is typically joined by a second shaft to a fan disposed upstream from the compressor in a typical turbo fan aircraft engine configuration for powering an aircraft in flight.
As the combustion gasses flow downstream through the turbine stages, energy is extracted therefrom and the pressure of the combustion gas is reduced. The combustion gas is used to power the compressor as well as a turbine output shaft for power and marine use or provide thrust in aviation usage. In this manner, fuel energy is converted to mechanical energy of the rotating shaft to power the compressor and supply compressed air needed to continue the process.
In the prior art, a known means for increasing performance of a turbine engine is to increase the operating temperature of the engine, which allows for hotter combustion gas and increased extraction of energy. Therefore, during operation and due to extreme temperatures of the combustion gas flow path and operating parameters, the stator blades may become highly stressed with extreme mechanical and thermal loading. However, until the creation of the present embodiments, the prior art has been unable to provide apparatuses, methods and systems that are as effective as the embodiments herein in reducing operational stress and/or temperatures placed on stator nozzle segments, including stator vanes, thereby allowing them to better withstand higher temperature operating conditions.
With respect again to embodiments herein, some engines include the stator nozzle assembly being, for example, an annular ring formed as a single piece. Other engines include the turbine stator nozzle assembly being formed as an annulus by a plurality of stator nozzle segments arranged in an annular array. The nozzle segments each include an inner band, an outer band and a vane extending therebetween. The vanes are hollow and receive a portion of pressurized air from the compressor which is used for cooling the vanes at all times during operation but especially during extremely high operating temperatures, specifically under certain conditions such as take-off or steep climbs. Additionally, compressed air may be released through surface apertures in the nozzle vane to form a thermal barrier of relatively cool air around the vane also called a cooling film.
In the past, an unmet need has been that, due to curvature near the leading edge of the vane, the compressed air apertures cannot be adequately provided to allow for even higher temperature operation. Additionally, due to vane curvatures in this area, the compressed air apertures cannot be created per intent successfully and thus the cooling film will not always adequately “attach” to the vane surface leaving the vane exposed to the high temperature combustion gas flow path.
As may be seen in this background section, there is a need for increasing durability of nozzle vanes which allow for increased operating temperatures and therefore increased turbine performance. It is further desirable to extend the useful service life of the nozzle. The present embodiments meet such needs and desires, and more.